Automatic approach pitch axis control system for aircraft

ABSTRACT

The invention relates to a system utilizing an exponential control law for glide slope capture. The capture maneuver from above or below the beam, is a function of decreasing glide slope beam error. The present autopilot approach coupler is an altitude rate command system which provides switchless signal processing during glide slope capture, and tracking.

This is a continuation of application Ser. No. 342,343 filed Mar. 19,1973, now abandoned. This last mentioned application is a division ofapplication Ser. No. 221,958 filed Jan. 31, 1972, now U.S. Pat. No.3,801,049 and relates to signal processing for automatic approach of anaircraft and more particularly relates to an improved system forautomatically controlling the pitch axis of an aircraft during anautomatic approach.

Prior art control systems which fly an aircraft close to the landingrunway and thereafter approach that runway and flare-out for touchdownare available as exemplified by U.S. Pat. No. 3,327,973 to KRAMER ET AL.However, such a system utilizes a landing control law which requires theflight crew to preselect a reference flight path approach by theaircraft to the runway which reference flight path may or may not beoptimized for the desired flight path. During normal flight and prior toapproach for automatic landing, the system in the above referencedKRAMER ET AL patent utilizes an automatic system for controlling theelevators and thus the pitch attitude of the aircraft. Engagement of theautomatic landing system with preselected reference flight path in theaforementioned manner by the flight crew which results in less thanoptimum flight path acquisition further results in abrupt movements ofthe aircraft and large initial flight path errors. Such abrupt movementsare highly objectionable in commercial aviation since causing passengeralarm and discomfort, and also very importantly, large initial flightpath errors limit the ability of system performance at low altituderesulting in consequent deterioration of noise abatement procedures.

It is accordingly an object of the present invention to provide a pitchaxis control system for providing smooth and accurate acquisition of theglide slope beam to prevent large errors in flight path at extremely lowaltitudes.

It is a further object of this invention to provide a pitch axis controlsystem permitting capture of glide slope error independent of glideslope angle and approach speed for various altitudes of glide slopecapture.

It is yet a further object of the present invention to provide in anautopilot control system, a synchronized automatic exponential captureof the glide slope error independent of glide slope angle.

The above and further objects are achieved in the present invention bysignal processing means for coupling control signals to utilizationmeans, e.g., the pitch axis control system which processing meansemploys a single set of control laws for signal processing during glideslope capture and tracking.

Other objects, features and advantages of the present invention willbecome apparent from the following description read on the accompanyingdrawings, wherein:

FIG. 1 corresponds to FIG. 1 of U.S. Pat. No. 3,327,973 which isincluded for ready reference to assist in comparison of the system ofthe present invention with this prior art;

FIG. 2 is a diagram useful in showing geometric relationships of theaircraft in relation to ground for developing the equations of flightpath control of the system of the present invention;

FIG. 3 is a block diagram showing signal processing utilized to deriveaircraft path command signals in accordance with the system of thepresent invention deemed helpful in further development of the equationsof flight path control of the system;

FIG. 4 is a graph showing commanded flight path as a function of a timesubsequent to glide slope capture further helpful in understanding thederived equation for commanded flight path;

FIG. 5 is a block diagram representative of signal processing foraircraft short period damping satisfied in the system of the presentinvention;

FIG. 6 is a block diagram illustrative of signal processing of thepresent system during the flare command phase of landing of theaircraft;

FIG. 7 is a block diagram showing system signal processing during thego-around phase;

FIG. 8 is a block diagram similar to FIG. 7 however in a more detailedaircraft environment;

FIG. 9 is an exemplary circuit embodiment of the go-around systems ofFIGS. 7 and 8;

FIG. 10 is a block diagrammatic representation of an embodiment of apitch axis control system according to the present invention; and

FIG. 11 shows in more detail the acceleration normal to flight pathdetector of FIG. 8;

FIG. 12 is an actual plot showing actual flight path compared to desiredflight path illustrative of the present pitch axis control systemperformance during glide slope capture tracking and go-around; and

FIG. 13 is an actual plot showing pitch axis control system performanceincluding flare, touchdown and automatic noise lowering.

Turning now to the system of FIG. 1 which is representative of the priorart, a comparison therewith will be made with the system of FIG. 10which is illustrative of the system of the present invention to bringout the features of the present system. The features of the presentsystem may then become focussed upon and appreciated in the subsequentanalysis from a signal processing standpoint and later system embodimentdescription which further explain and amplify how these features andresultant advantages are achieved in accordance with the present pitchaxis control system.

The system of FIG. 1 provides a synchronized automatic capture of theglide slope error which depends upon pilot initiated computer inputinformation based upon glide slope angle, approach speed, etc., which isonly optimized for one set of environmental or airplane conditions andfor one capture altitude while the present system of FIG. 10 provides asynchronized automatic exponential capture of the glide slope errorindependent of glide slope angle, approach speed, wind conditions, etc.,which allows optimum performance under substantially any altitude ofglide slope capture.

The system of FIG. 1 utilizes pitch altitude for minor loop stabilityresulting in looser control of the desired flight path in the presenceof wind and tends towards less reliability due to the added sensor whilethe present system of FIG. 10 does not require the use of a pitchattitude source for minor loop stability thus allowing for more accuratecontrol along the desired flight path and also eliminating the extrasensor reliability.

The system of FIG. 1 can be seen to utilize a vertical velocity computerto derive velocity errors relative to true vertical, not the desiredflight path. The system also requires a longitudinal accelerometer foroptimum compensation for wind conditions or airplane speed bleeds. Thepresent system of FIG. 10 in contrast uses a normal accelerometerhowever tilted relative to the aircraft body axis (see FIG. 11 for moredetail) to provide instantaneous normal velocity errors relative to thedesired flight path and to compensate for any longitudinal accelerationerrors due to environmental conditions or aircraft speed bleeds. Thepresent system further eliminates the need for a longitudinalaccelerometer to compensate for these errors.

The system of FIG. 1 utilizes a vertical velocity computer whichcomputes the vertical velocity and does not contain pitch rateinformation requiring both pitch rate and pitch attitude for minor loopstability, and further requires the monitoring of these sensors forautomatic landings. The present system uses a normal accelerometermounted forward of the aircraft center of gravity to provide a signalproportional to pitch acceleration which is passed through a lag filterto provide a pitch rate signal. The present system thus eliminates thepitch rate gyro as a critical sensor thus facilitating easier monitoringof the system.

The system of FIG. 1 utilizes a fixed vertical beam sensor switch pointdetector which is optimized for only one glide slope capture altitudeand is much less acceptable for lower altitude glide slope captureswhile the present system of FIG. 10 utilizes a vertical beam sensor(switch point detector) that is downstream of the glide slope gainprogrammer. This allows optimum glide slope capture at substantially anyaltitude by varying the glide slope capture point inversely withaltitude which is advantageous in noise abatement type approaches.

The present system embodiment of FIG. 10 provides an automatic go-aroundcommand as does the system of FIG. 1, however the system of FIG. 10utilizes the same circuitry already utilized in FIG. 10 to perform otherILS coupling functions and has the following features and functionaladvantages over the system of FIG. 1:

a. If go-around circuitry fails, the system of FIG. 10 will flare theairplane allowing time for pilot correction at extremely low altitudes.

b. In the present system, the initial go-around command is independentof the final go-around command thereby allowing the aircraft to initiatego-around and assume positive rate of climb even if final command hasfailed.

c. The flare command is not inhibited by go-around which additionallyreduces altitude loss at extremely low altitudes, and allows automaticgo-arounds even after the aircraft touches down.

d. The present system circuit design is such that no failure of thego-around command can result in an increased sink rate of the airplaneafter initiation of go-around.

The system of FIG. 1 does not provide automatic noise lowering aftertouchdown. The pilot is required to disengage the autopilot aftertouchdown and lower the nose to ground manually prior to braking theaircraft while the present system of FIG. 10 provides automatic noselowering after touchdown which allows the pilot to leave the systemengaged after touchdown and puts in nose down elevator to hold theaircraft on the ground after touchdown.

The present pitch axis system circuit embodiment implements thefollowing 4 control equations:

1. The command path functions ##EQU1##

2. The aircraft short period damping function

    S.sub.s =S.sub.s.sbsb.o +τ.sub.2 K.sub.2 (Vn.sub.o -Vn)+τ.sub.2 K.sub.2 <(θ-θ)+K.sub.3 (θ.sub.o -θ)

3. The flare command function ##EQU2##

4. The go-around function for zero h° error ##EQU3##

FIG. 2 showing aircraft relative position, viz., geometricrepresentation of glide slope geometry and FIG. 3 showing in block formaircraft path command signal processing may now be considered indeveloping the path error and then path command signal terms of thepresent pitch axis control system where

z = angular error of airplane from glide slope center as sensed by glideslope error detector

θ₁ = glide slope center reference angle

θ₂ = θ₁ - z

h = distance from glide slope center and airplane receiving antennaperpendicular to glide slope center

h₁ = distance from airplane receiving antenna and ground perpendicularto glide slope center

h₂ = distance from glide slope center and ground perpendicular to glideslope center

h₃ = vertical distance from glide slope center and ground ##EQU4##

θ₁ -θ₂ = z ##EQU5## But ##EQU6##

    Δx = h.sub.3 TANθ.sub.1

and

    h.sub.3 =  g(x) ##EQU7##

    h = KvZ ##EQU8##

For zero path error, the path commanded relative to the glide slope zeroplane is defined by: ##EQU9##

This is in the form: ##EQU10## where: ##EQU11## and z = damping; w_(z) =natural frequency hence: ##EQU12## Then h(s) is in the form: ##EQU13##

The above path command equation h.sub.(t) in terms of time constants,natural frequency and damping is seen to result in a glide slope capturewhich is always exponential (see FIG. 4) and which is always enteredtangentially determined by I_(o).

Turning now to the position of the system providing short period dampingshown in FIG. 5 the surface command equations are developed in thefollowing:

    ______________________________________                                           surface .sub.t.sub.=0 =(V.sub.N0 +Lθ )K.sub.2 τ.sub.2            +θ.sub.0 K.sub.3 +S.sub.S0 -I.sub.0.sbsb.2 =0                               I.sub.0.sbsb.2 =V.sub.N.sbsb.0 K.sub.2 τ.sub.2 +Lθ K.sub.2      τ.sub.2 +θ.sub.0 K.sub.3 + S.sbsb.0                                    surface .sub.t.sub.=.sub.+0 =V.sub.N K.sub.2τ.sub.2 +LθK.sub.    2τ.sub.2 +θK.sub.3 +S.sub.S -I.sub.0.sbsb.2                           surface= for airplane damping satisified:                                   0=  surface .sub.t.sub.=10=τ.sub.2 K.sub.2 (V.sub.N -V.sub.N0)+           τ.sub.2 K.sub.2 L(θ -θ.sub.0)+K.sub.3 (θ-θ.sub    .b + .sub.S - .sub.S0                                                          .sub.S = .sub.S0 +τ.sub.2 K.sub.2 (V.sub.N0 -V.sub.N)+ τ.sub.2       K.sub.2 L(θ .sub.0 -θ)+K.sub.3 (θ.sub.0 -θ            ______________________________________                                    

Normally at glide slope capture, the terms VN_(o), θ_(o) and θ_(o) arevery small and can be neglected therefore:

    Ls=Ls.sub.o +τ.sub.2 K.sub.2 VN + τ.sub.2 K.sub.2 L θ+K.sub.3 θ

for flare command, the signal processing elements of the system areshown in FIG. 6 which results in flare command signals derived asfollows:

    h°ERROR=h°+h°.sub.o +h° COMMAND SINCE Kv=0

Where ##EQU14##

And for zero path error: ##EQU15##

Turning now to FIG. 7, the following derivations show how go-aroundequations representative of these signals are developed by the positionof the system shown in FIG. 7: ##EQU16##

While the system portion shown in FIG. 7 and the above equations definethe go-around command signals generated the explanation which follows inconnection with FIGS. 8 and 9 will further serve to explain in a morephysical sense and in a complete circuit schematic respectively how thego-around function is achieved in the aircraft environment.

From the preceding block diagram and discussion, it should be noted thatthe automatic go-around command used in the present autopilot approachsystem is not automatically initiated but requires pilot activation ofthe go-around switch of FIG. 8 (correspondingly switching means 24 ofFIG. 7 comprising a transistor switch) which is preferably located onthe throttle levers. If during an approach, the flight crew decides thatconditions are not adequate to continue the approach, e.g., traffic onthe runway, or inadequate visibility for landing, the pilot can initiatean automatic go-around by increasing the thrust and activating thego-around switch. This action will cause the autopilot to command theairplane to fly a programmed rate of climb. The programmed rate of climbis generated in the form of first and second signal components asfollows:

1. The first signal component generated in the go-around portion of thesystem is shown in FIG. 8 commands a climb rate of 300 feet per minute.The h° error signal which is proportional to elevator command, can beseen to be composed of three terms prior to go-around which are:

Prior to flare

1. g/s displacement (short term h° command) programmed to zero at 65feet

2. G/S integral (h° command reference)

3. h° response (damping terms)

During flare

1. flare command (held at zero output until an altitude of approximately53 feet)

2. G/S integral (h° command reference)

3. h° response (damping terms)

2. The second signal component generated by the go-around portion of thesystem shown in FIG. 7 is the term which actually causes the aircraft toperform an automatic go-around. This term (h° command reference) isproportional to the aircraft rate of sink when the aircraft isconducting an approach since when the aircraft is flying zero glideslope error (on glide slope centerline) the output of the integratorcircuit 9 must be equal and opposite to the h° response of the dampingterm of lag filter 17 to null out the h° error and fly a zero elevatorcommand. This h° command reference is a fly down command so that theaircraft is descending at approximately 600 to 700 feet per minute onthe centerline of the glide slope. When the pilot initiates an automaticgo-around by pushing the go-around switch 24 (see FIG. 8), two eventsoccur: first, the glide slope displacement and integral input paths areremoved by switch S_(A) so that no reference to the glide slopecenterline is maintained during the go-around. This in itself does notcause any go-around command to be generated but causes the circuitry tomaintain an h° hold command (a fixed output on the integrator circuitsince the input to the integrator is zero) prior to flare or if in flare(at less than an altitude of about 53 feet) to continue to flare theaircraft due to the flare command. The second event occurssimultaneously with the closing of a resistive circuit path 142 viaswitch S_(B) around the integrator circuit which washes out the glideslope integrator generated signal, h° command reference. Since theoutput of the integrator circuit is a fly down command, washing out orelimination of the integrator output signal is representative of a flyup command having a time constant determined by the RC network formed bythe switched resistive circuit path and the capacitor providing theintegrator feedback. For a Boeing Airplane Company type 747 aircraft,this time constant equals approximately 4.5 seconds but is dependentupon the particular aircraft characteristics. This function causes theaircraft to break its rate of sink. In addition as can be seen in FIG.8, a voltage bias is summed in through a resistive network (not shown)to cause the aircraft to initially seek a climb rate of 300 feet perminute.

A second phase in go-around occurs after closing of switches S_(A) andS_(B) when the aircraft's flaps are raised to less than 23° to providethe go-around flap setting thereby switching in an additional gain pathfrom the voltage bias and causing the lag circuit to command anadditional 700 feet per minute climb rate for a total command rate of1000 feet per minute.

An actual exemplary embodiment of the go-around system of FIGS. 7 and 8is shown in FIG. 9. If the go-around is initiated below 53 feet (flareregion; or just prior to 53 feet and the aircraft enters the flareregion, the flare computer will also command a decrease in rate of sinkwhich aids the go-around command and allows the automatic go-around tobe used safely at very low altitudes including after touchdown.Automatic nose lowering after touchdown is provided in the present pitchaxis control system which does not utilize pitch attitude as a dampingterm but in which primary damping is dependent on altitude rate. Attouchdown, the flare command is requesting a sink rate of 2.5 to 3 feetper second. When the aircraft lands, the aircraft sink rate is reducedto zero in a very short time interval which produces an error betweenthe actual sink rate of zero and the commanded sink rate of 2.5 to 3feet per second. This results in a nose down elevator command effort forproviding an increase in sink rate to 2.5 to 3 feet per second. Thepresent pitch axis control system damping permits this maneuver in acontrolled manner. Prior art systems which utilize pitch attitude fordamping cannot generate sufficient altitude rate error at touchdown tolower the nose, hence the pilot must disconnect the control system andlower the nose manually.

Turning now to FIG. 10, there is shown the complete control system whichprovides the several functions, e.g., flare command, go-around, etc.,already separately discussed. In the following discussion referencenumerals corresponding to those used earlier will be used to identifycorresponding elements of the system.

In the system of FIG. 10, between the system output terminal 15 and thesumming junction 10 there is coupled a negative feedback loop. Thisnegative feedback loop comprises the glide slope integrator 9 connectedthrough switching means 12 comprising a relay switch in the positionshown 2ND series gain 140 for providing a synchronizing path. Thissynchronizing path provides two functions when operating in thesynchronizing mode. The first function is for reducing signals presentat the system output terminal 15 to reference potential (zero) bydriving glide slope integrator circuit 9 so that the output signalvoltage of integrator circuit 9 is substantially equal and opposite tothe sum of the remaining signal voltages at summing junction 10. in thismanner, the pitch axis control system output signal at computer systemoutput terminal 15 is maintained at reference potential (zero voltagelevel) to assure that no undesirable aircraft maneuver is experienced atthe time of engagement of the automatic approach and landing computer ofthe present pitch axis control system. The second function of thesynchronizing path is for providing glide slope capture initialconditions so that when the present automatic approach and landing pitchaxis control system is engaged by closing switching means 12 to thedotted line position, the present system will maneuver the aircraft ontothe glide slope zero plane. This function is accomplished in a uniqueand novel manner without having to switch in a separate signalgenerating means and by utilizing the same control laws previouslyderived which also provide the glide slope zero plane tracking. Sincethe glide slope integrator circuit 9 has stored at its output, a signalvoltage which is equal and opposite to the sum of all other signalvoltages appearing at the input of summing junction 10, and, for a glideslope capture from a point below the glide slope zero plane, this storedoutput contains a signal which is equal and opposite to the fly upcommand from the glide slope error detector 4 through the variable glideslope gain programmer circuit 11. Circuit 11 comprises means well-knownin the art for multiplying two variables together e.g., pulse widthmodulated shunt switching means. At a fixed error signal level from theglide slope zero plane, the vertical beam sensor 66, threshold isexceeded causing switching means 12 to transfer and thus removing theoutput signal at terminal 15 from the input summing junction 8 of theglide slope integrator. The synchronizing path is removed by this actionand the glide slope integrator signal at this instant in time is fixedand can no longer change to drive the output signal at terminal 15 tozero for any change in the remaining input signals to summing junction10. As the aircraft continues to fly toward the glide slope zero plane(see FIG. 2), the fly up command from glide slope error detector 4 isreduced in magnitude thus creating an error signal at system outputterminal 15 in a fly down command direction which comprises the storedfly down signal from glide slope integrator 9 and the decreased fly upsignal from glide slope error detector 4. The fly down command errorsignal at the output terminal 15 causes the elevator control system tocause displacement in a direction causing the aircraft to descend. Thisdisplacement of elevator surfaces in the control system is coupled bysurface feedback measuring means 16 to null the system output commandsignal voltage at system output terminal 15. The aircraft rate ofdescent signal voltage provided by altitude rate detector 2 and theaircraft rate of acceleration signal voltage provided by detector means1 (comprising an accelerometer having sensitive axis mounted normal tothe desired flight path) shown in FIG. 10 (and in complete detail inFIG. 11) sense that the aircraft is descending and these two signalvoltages are summed and coupled through lag filter means 17 (comprisinga low pass lag filter, e.g., a resistor in parallel with a capacitor infeedback circuit of an operatinal amplifier) to produce a signal whichis referenced to the aircraft flight path for short period maneuveringand to the aircraft vertical rate of descent for long term maneuvering.This uniquely derived signal is obtained by combining at adder 135(comprising, e.g., a summing junction): higher frequency signalcomponents from an accelerometer 1 which is tilted physically in theaircraft such that its sensing axis is disposed perpendicular to theflight path of the aircraft and which is positioned forward of thecenter of gravity of the aircraft (as shown in FIG. 11) which transmitsthese higher frequency signal components through high pass filtercircuit 131 and summing resistor 132, and lower frequency componentsfrom an altitude rate signal source 2 which is reference to verticalrate of descent. As the aircraft descends, the output signal from lagfilter 17 having the above higher and lower frequency signal componentsis representative of a fly down response in the system or a deviationfrom the aircraft flight path available through the circuit path coupledto junction 10 to null or cancel the signal voltage representative ofcommanded deviation from the aircraft flight path 119.

The results of the above described method of acquiring the glide slopezero plane is a fly down (or fly up if approaching from above the glideslope zero plane) altitude rate command signal voltage proportional tothe error between the stored glide slope error signal voltage at theoutput of glide slope integrator circuit 9 and the actual glide slopeerror signal voltage generated by glide slope error detector 4. In thismanner this unique feature of the present pitch axis control systemprovides a means for acquiring the glide slope zero plane which issubstantially independent of external factors such as aircraft speed,glide slope zero plane angles, and glide slope error signal gradients.The above feature is accomplished by utilizing only one signalgenerating means for both glide slope capture and tracking functions.

The above described pitch axis control system provides a flight pathcommand signal at the system output terminal 15 which positions theaircraft on a flight path to exponentially acquire the glide slope zeroplane, and it will be further noted that the closing of the switch 12(to the position shown by the dotted line) also couples in seriescircuit path glide slope gain programmer circuit 11 2ND gain 141 betweenglide slope error detector circuit 4 and summing junction 8 therebyproviding a means for varying the stored glide slope error signalpresent at the output of glide slope integrator 9 during the glide slopeacquisition maneuver and subsequent glide slope zero plane tracking tothereby eliminate errors developed in the flight path command signalpresent at system output terminal 15 and as a consequence cause theaircraft to acquire and track the zero plane of the glide slope errorsignal. The glide slope integrator output signal voltage from integratorcircuit 9 at this time is proportional to but of opposite polarity tothe descent rate signal voltage of the aircraft at low pass lag filter17 which relationship is required to maintain the glide slope errorsignal voltage at glide slope error detector 4 equal to zero.

The damping terms for the pitch axis control system of FIG. 10 arederived in a novel and unique manner by mounting of the accelerometer 1in the manner shown in FIG. 11, viz., normal to the flight path andforward of the aircraft center of rotation such that the output ofaccelerometer circuit 1 comprises: a signal voltage component ##EQU17##proportional to the time rate of change of the aircraft velocity normalto the desired flight path; and voltage component ##EQU18## proportionalof the rate of change of aircraft pitch attitude rate, and which is alsoinsensitive to the time rate of change of the aircraft's velocitytangential to the flight path ##EQU19## An additional `versine` signal136 derived in a manner well-known in the state of the art is generatedfrom roll angle sensor 140 to compensate the accelerometer sensor 139and eliminate the effects of the versine term ##EQU20## inherent in abody mounted accelerometer. The output signal voltage from accelerometercircuit 1 is processed through lag filter 17 which provides an outputsignal voltage which is proportional to time rate of change of aircraftpitch attitude ##EQU21## and velocity normal to flight path, VN. Asecond circuit path is provided in series circuit between accelerometer1 and system output terminal 15 by means of lag filter 18 connected inparallel with lag filter 17 to provide a further output signal voltagewhich is proportional to the output of accelerometer 1 normal to flightpath. In this manner, the critical damping terms necessary for stabilityof the aircraft when flying an approach and landing are derived from thesingle source (accelerometer 1) thereby increasing the reliability ofthe system by reducing the number of critical component sensorsnecessary in the achievement of safe stability margins.

The pitch rate voltage signal source utilized is pitch rate detector 3which is coupled in series circuit through band pass filters 111 andsumming resistor 113 to summing junction 10 to provide an additionaldamping term in the system output signal voltage at output terminal 15by summation through summing junction 10. This damping term is notcritical in affecting aircraft or flight path stability.

A further feature of the presently described pitch axis control systemof FIG. 10 provides a unique and novel means of allowing glide slopeacquisition maneuvers at substantially any distance from the landingrunaway (or from substantially any altitude above the runway). Thisaspect of the system hereinafter described has important considerationsin connection with noise abatement approaches wherein it is desirable toacquire the glide slope zero plane as close to the desired landing pointon the runway as possible to avoid long approaches over populated areas.In this respect, the system embodiment of FIG. 10 utilizes a verticalbeam sensing means 66 (e.g., a threshold detector) which is locateddownstream in terms of signal processing from the gain programmercircuit 11. The gain programmer circuit 11 varies the glide slope errorsignal gain path (which includes the coupling of glide slope errordetector 4, gain programmer circuit 11, lag filter means 115 (comprisinga low pass filter) and summing resistor 117 in series circuit path tosumming junction 119) to convert the angular glide slope error signalvoltage from glide slope detector 4 into an error signal voltage whichis proportional to distance from the glide slope zero plane. In thismanner, the response of the system of FIG. 10 to glide slope errors ismaintained constant at substantially any altitude down to that altitudelevel at which the programmer output signal voltage from programmercircuit 11 is programmed to zero immediately prior to flaring of theaircraft. Since vertical beam sensor 66 is coupled in circuit betweenthe glide slope gain programmer circuit 11 and the system outputterminal 15 (downstream of the glide slope gain programmer circuit 11 interms of signal processing) the present system of FIG. 10 can maintain asubstantially constant distance from the glide slope zero plane forsystem activation irrespective of the distance from the runway that thesystem is engaged. This means that for low altitude glide slopeacquistions, the vertical beam sensor 66 detection threshold is exceededfor greater error output signal voltages from glide slope error detector4 than it does for higher altitude glide slope acquisitions whichresults in an aircraft maneuver and flight path performance which issubstantially identical for both high and low altitude captures. Theunique and novel feature allows the present automatic approach andlanding pitch axis control system to be utilized in the above manner fornoise abatement approaches heretofore not possible.

A unique and novel flare command is provided by the present system whichis switchless and provides tighter control of landing dispersions alongthe runway. The flare command signal voltage at circuit connection 19 inthe series circuit comprising: altitude above terrain detector 5connecting through limiter circuit 6 (comprising voltage limiting means,e.g., a saturated amplifier), the parallel combination of altitude ratecircuit 21 comprising a high pass filter and altitude path displacementcircuit 20 comprising a summing resistor proportional to displacementgain to summing junction 7, asymmetrical limiter circuit 125, circuitconnection 19, summing resistor 127 to adder 129, adder 119, summingjunction 10, and amplifier means to system output terminals 15, is aflare command having a flare point and a touchdown rate of descentcommand which are varied automatically by mechanization of the controllaws to provide tight control of the aircraft landing dispersions due toenvironmental conditions such as winds, terrain and varying aircraftflight parameter such as gross weight, center of gravity, flapconfiguration, and speed. The output voltage of the altitude aboveterrain detector 5 is limited by limiter circuit 6 at an altitude sothat large irregularities in terrain distant from the normal flare pointof the aircraft approaching the landing runway do not affect the flarecomputation. Above the predetermined altitude for which the limiter isset, the output of limiter circuit 6 is a fixed parameter, i.e., notvarying with time. The output voltage from altitude rate circuit 21 iszero since the input voltage to circuit 12 is not time varying, and theinput voltages to summing junction 7 comprose the predetermined leveloutput voltage from voltage limiter circuit means 6 coupled throughaltitude path displacement circuit 20 which provides displacement pathvoltage amplification and the zero voltage output of altitude ratecircuit 21. The input of the flare command signal voltage at lead 22transmitted to summing junction 10 is limited by asymmetrical limitercircuit 125 such that for positive summation of the input voltages tosumming junction 7, no change in the output level of limiter circuit 125can occur.

As the aircraft descends below the altitude at which the input voltageto limiter circuit 6 causes saturation the output voltage of limitercircuit 6 decreases in a manner proportional to altitude above theterrain. The output voltage from rate circuit 21 senses the rate ofchange of altitude with time and when the sum of the output voltages ofrate circuit 21 and altitude displacement circuit 20 coupled intosumming junction 7 is negative in polarity, the output voltage 19 fromasymmetrical limiter circuit 125 is a command signal voltage on lead 22coupled to summing junction 10 representative of a decreasing altituderate command. The preceding circuit feature enables an aircraft which isdescending at a high sink rate to being to command a flare maneuversooner than an aircraft descending at a lower sink rate. As the aircraftenters the flare region and approaches touchdown, heretofore, wind gustsor misapplication of thrust by the pilot have caused the aircraft to"float" however in accordance with the above discussed features of thepresent system, the flare command touchdown sink rate is caused to varyas a function of time in order to reduce increased landing dispersionnormally experienced under the above and other conditions. As theaircraft begins to decrease its sink rate prior to touchdown, the ratecircuit 21 output voltage begins to decrease. If the aircraft begins to"float" (i.e., approaches zero sink rate) at some altitude above therunway, the summed output voltage from summing junction 7 decreases dueto the decreasing voltage from rate output circuit 21 hence decreasingflare command called for by the flare command signal voltage on lead 22thereby causing the aircraft to increase its rate of sink for reducingtouchdown dispersion.

The present system control laws discussed earlier as implemented in thepresent system embodiment allow generation of a switchless flare commandnot susceptible to switch failure prohibiting flare and further allowtouchdown dispersions due to environmental and aircraft parameters to beminimized in the manner hereinbefore discussed.

In addition to the preceding, the present system includes circuitfeatures which generates an automatic go-around command signal voltageat the output of adder 23 which is generated in a manner such that asthe aircraft enters the flare region an additional go-around commandsignal voltage is generated as a portion of the switchless flare commandsignal voltage present on lead 22 to reduce the altitude loss during thego-around maneuver. The features of the go-around circuitry which areunique in the present system are that all of the same circuit componentsutilized conducting the approach are utilized which are already knownoperative prior to initiation of go-around. Activation of the go-around(G/A) switch 24 by the pilot causes switch 14 to move the open position(shown by dotted line) thereby reducing the output of the glide slopegain programmer circuit 11 to zero and further causing switch 13 toclose (shown by the dotted line). Closing switch 13 connects togethersumming junctions 8 and 23 which results in conversion of glide slopeintegrator circuit 9 into a lag circuit through gain 142 with timeconstant γ= 1/K₈. Since the output voltage from the glide slopeintegrator circuit 9 is proportional to the actual descent rate of theaircraft and is representative of a fly down command, the closing ofswitch 13 causes subsequent "washout" elimination in the output ofintegrator circuit to a resultant zero fly down command thereby causingan error signal to be generated at summing junction 10 to cause theaircraft to break its sink rate and assume level flight. If the aircraftis in the flare region, the flare command signal voltage present on lead22 will command the aircraft to climb to an altitude equivalent to thepreviously referred to flare initiation altitude and then maintain thataltitude. These unique circuit features provide a go-around maneuverwhich is fail safe in the sense that it uses "known to be operating"components and does not require the introduction of a second signalsource to initiate the go-around maneuver (as is the case in the FIG. 1system representative of the prior art), place the aircraft in levelflight and maintain an altitude above the flare region.

A predetermined voltage level is provided by go-around bias source 26which is summed into summing junction 23 thereby causing the outputsignal voltage of glide slope integrator circuit 9 to command a climbrate when switch 13 is closed.

Those skilled in the art will appreciate the important and significantfeature of the present system which provides a go-around command whichis fail safe in that it cannot inhibit a normal flare of the aircraft ifit fails and the further important feature that the system cannot causea nose down hardover situation as a result of failed components withinthe go-around circuitry and the further feature that the initialgo-around maneuver utilizes the same signal generating means alreadyknown to be operative.

FIG. 12 is a plot of actual airplane performance during a glide slopeacquistion, track and subsequent pilot initiated automatic go-around fora Boeing 747. As can be seen the principles of this invention providesmooth and accurate acquisition of the glide slope zero plane(represented by zero glide slope error on the plot) and accuratetracking of the glide slope zero plane, and smooth and accuratego-around maneuver.

FIG. 12 includes automatic landing. Again the glide slope trackingaccuracy prior to flare can be measured to be inches by those skilled inthe art. The automatic nose lowering feature after touchdown can be seento be smooth providing the pilot the opportunity to keep his attentionon the task of stopping the airplane while the autopilot performs thetask of keeping the airplane on the ground.

What is claimed is:
 1. In combination in a pitch axis control system fora multi-engine jet transport aircraft;first means for generating a firstsignal representative of command deviation from aircraft flight path;second means for producing a second signal insensitive to accelerationsalong the aircraft flight path and referenced to the aircraft flightpath for short period maneuvering and to the aircraft vertical rate ofdescent for long term maneuvering; third means comprising a lag filtercoupled to said second means for producing a third signal representativeof a fly up or fly down response in the system or a deviation from theflight path; fourth means for combining said first signal and said thirdsignal to null or cancel said first signal; and fifth means includinglongitudinal control actuator means coupled between said fourth meansand the control surfaces of said aircraft; and wherein said second meanscomprises: an accelerometer which is tilted physically in said aircraftso that the sensing axis of said accelerometer is disposed perpendicularto the flight path of the aircraft, said accelerometer also positionedforward of the center of gravity of the aircraft; an altitude ratesignal source; and, means for combining the outputs of saidaccelerometer and altitude rate signal source to provide said secondsignal.
 2. In combination in a pitch axis control system for providing aflight path command signal at an output terminal:glide slope errordetector means; gain programmer circuit means, amplifier means and lagfilter means coupled in series circuit between said glide slope errordetector means and said output terminal; a negative feedback circuitpath including gluide slope integrator circuit means coupled betweensaid output terminal and the input terminal of said amplifier means;and, vertical beam sensing means coupled in said series circuit betweensaid gain programmer circuit and said amplifier for interrupting saidnegative feedback circuit path and coupling said glide slope gainprogrammer circuit to said glide slope integrator circuit means.
 3. Theinvention according claim 2 wherein said vertical beam sensing means isinversely programmed by said gain programmer circuit means providingglide slope capture threshold which varies glide slope capture pointinversely with altitude to provide a glide slope capture maneuver whichis independent of altitude above terrain.
 4. In combination in a pitchaxis control system for causing an aircraft to acquire a glide slopezero planefirst means for producing a first signal representative ofangular deviation from said glide slope zero plane; second meansincluding gain multiplier means responsive to said first signal forproducing a second signal representative of linear distance from saidglide slope zero plane; and, third means including sensor means fordetecting a predetermined linear distance from said glide slope zeroplane thereby causing said pitch axis control system to acquire saidglide slope zero plane.
 5. In combination in a pitch axis control systemfor an aircraft to provide an error signal at an output terminalfirstmeans for generating a first signal representative of the deviation ofsaid aircraft from a desired flight path or glide slope zero plane;second means for producing a second signal representative of thevelocity of said aircraft normal to said flight path for short termmaneuvering and to the vertical velocity of said aircraft for long termmaneuvering of said aircraft; third means for generating a third signalrepresentative of commanded vertical velocity of said aircraft tomaintain said desired flight path or said glide slope zero plane; and,fourth means for combining said first signal, said second signal, andsaid third signal to null or cancel said first signal to provide saiderror signal at said output terminal.
 6. The invention according toclaim 5 wherein said output terminal is coupled to fifth means includinglongitudinal actuator means for controlling the longitudinal controlsurfaces of said aircraft.
 7. In a pitch axis control system for anaircraft first means for producing a first signal representative of theacceleration of the aircraft normal to a glide slope zero plane;and,second means including washout circuit means and lag circuit meanscoupled in series circuit path with said first means for producing asecond signal representative of the instantaneous velocity of saidaircraft normal to said glide slope zero plane.
 8. The inventionaccording to claim 7 wherein said first signal comprises the outputsignal of an accelerometer disposed in said aircraft such that thesensing axis of said accelerometer is perpendicular to said glide slopezero plane.
 9. Apparatus for guiding a craft to a landing along a radiodefined glide slope beam, comprisingan integrator for storing a signalrepresentative of a desired descent rate the glide slope beam. means forproviding a signal representative of the craft displacement from thecenter line of the glide slope beam, mean for providing a signalrepresentative of the actual descent rate of the craft, means foralgebraically summing the stored signal of said integrator with theactual descent rate and beam displacement signals, and means forselectively coupling the output of said summing means to the input ofsaid integrator prior to the craft intercepting a predetermined radialof the glide slope beam for controlling the output signal level of theintegrator in accordance with the instantaneous amplitude of the beamdisplacement and actual descent rate signals whereby upon the craftintercepting said predetermined radial at a given velocity in a levelflight condition a desired descent rate signal is produced at the outputof said integrator to cause pitch down of the craft for capturing theglide slope beam.
 10. The apparatus of claim 9 wherein the selectivecoupling means is operative upon disconnecting the output of saidsumming means from the input of said integrator for connecting thedisplacement signal means to the integrator input so that the integratoroutput is driven to a level equal to that of the steady state actualdescent rate signal whereby the displacement signal provides thepredominant control for guiding the craft along the center line of theglide slope beam.
 11. An integrated Autopilot/Flight Director glideslope coupler system for guiding a craft to a landing along a radiodefined glide slope beam, comprisingmeans for providing a signalrepresentative of the craft displacement from the center line of thebeam. switching means enabling said system to be set initially in asynchronizing mode and thereafter in either an Autopilot or FlightDirector mode, means for providing a signal representative of the actualdescent rate of the craft, an integrator for storing a signalrepresentative of a desired descent rate, means for algebraicallysumming the integrator output signal with the displacement and actualaltitude rate signals, and said switching means being operative in afirst condition corresponding to the synchronizing mode, prior to thecraft intercepting a predetermined radial of the glide slope beam, toconnect the output of said summing means to the input of said integratorfor controlling the output signal level thereof in accordance with theinstantaneous amplitude of the beam displacement and actual descent ratesignals whereby upon the craft intercepting said predetermined radial ina level flight condition at a prescribed velocity the signal at theoutput of said integrator is representative of a desired descent rate tocause a craft pitch down to fly along the center line of the glide slopebeam.
 12. The apparatus of claim 11 including further means wherein theswitching means is operative upon completion of the synchronizing modeto disconnect the output of the algebraic summing means from theintegrator input and to connect said further means thereto.
 13. Theapparatus of claim 12 including means for continuously modifying thedisplacement signal in accordance with the instantaneous altitude of thecraft from a first predetermined altitude down to a second predeterminedaltitude to compensate for convergence of the glide slope beam.
 14. In apitch axis control system for providing a flight path command signal atan output terminal (15);first means coupled to said output terminal(15), said first means comprising second means (139) for generating asignal voltage having a first component proportional to the time rate ofchange of aircraft velocity normal to the desired flight path and asecond component proportional to the rate of change of pitch attituderate, and third means (140) comprising a roll angle sensor forcompensating said second means, said system further comprising fourthmeans coupled to said output terminal for providing a further dampingterm in the flight path command signal at said output terminal, saidfourth means comprising a pitch rate detector (3) and a band pass filter(111) coupled between said pitch rate detector (3) and said outputterminal (15).
 15. In a pitch axis control system for providing a flightpath error signal at an output terminal:first means for generating afirst signal (113), said first means comprising a pitch rate detector(3) and a series coupled band pass filter (111); second means forproviding a second signal (17), said second signal having a firstcomponent proportional to the time rate of change of aircraft velocityto normal to the desired flight path and a second component proportionalto the rate of change of said aircraft pitch attitude rate; third meanscoupled between said second means and asid output terminal to provide athird signal (18) having a third component proportional to said aircraftvelocity normal to said desired flight path and a fourth componentproportional to said aircraft pitch attitude rate; fourth means (10) forcombining said first signal, said second signal and said third signal toprovide said flight path error signal at said output terminal.
 16. Anintegrated glide slope couple system for guiding a craft to a landingalong a radio defined glide slope beam, comprisingmeans for providing asignal representative of the craft displacement from the center line ofthe beam, switching means enabling said system to be set initially in asynchronizing mode and thereafter in a glide slope mode, means forproviding a signal representative of the actual descent rate of thecraft, an integrator for storing a signal representative of a desireddescent rate, means for algebraically summing the integrator outputsignal with the displacement and actual altitude rate signals, and saidswitching means being operative in a first condition corresponding tothe synchronizing mode, prior to the craft intercepting a predeterminedradial of the glide slope beam, to connect the output of said summingmeans to the input of said integrator for controlling the output signallevel thereof in accordance with the instantaneous amplitude of the beamdisplacement and actual descent rate signals whereby upon the craftintercepting said predetermined radial in a level flight condition at aprescribed velocity the signal at the output of said integrator isrepresentative of a desired descent rate to cause a craft pitch down tofly along the center line of the glide slope beam.
 17. The apparatus ofclaim 16 including further means wherein the switching means isoperative upon completion of the synchronizing mode to disconnect theoutput of the algebraic summing means from the integrator input and toconnect said further means thereto.
 18. The apparatus of claim 17including means for continuously modifying the displacement signal inaccordance with the instantaneous altitude of the craft from a firstpredetermined altitude down to a second predetermined altitude tocompensate for convergence of the glide slope beam.
 19. In combinationin a pitch axis control system for providing a flight path commandsignal at an output terminal:glide slope error detector means; gainprogrammer circuit means, amplifier means and lag filter means coupledin series circuit between said glide slope error detector means and saidoutput terminal; a negative feedback circuit path including glide slopeintegrator circuit means coupled between said output terminal and theinput terminal of said amplifier means; and, vertical beam sensing meanscoupled in said series circuit between said gain programmer circuit andsaid amplifier.